Hourglass airfoil cooling configuration

ABSTRACT

A core structure for a providing a cooling passage in a gas turbine engine includes a core body that has a first passage core. The first passage core has a first width in a chord-wise direction near a first wall. A second width in the chord-wise direction near a second wall. A third width in the chord-wise direction between the first and second walls. The third width being smaller than the first and second widths to form an hourglass shape.

CROSS-REFERENCE TO RELATED APPLICATION

This application is a divisional application of U.S. application Ser.No. 16/515,528, filed on Jul. 18, 2019.

BACKGROUND

This disclosure relates to gas turbine engines and particularly tointernally cooled airfoils of rotor blades and stator vanes.

A gas turbine engine typically includes a fan section, a compressorsection, a combustor section and a turbine section. Air entering thecompressor section is compressed and delivered into the combustorsection where it is mixed with fuel and ignited to generate a high-speedexhaust gas flow. The high-speed exhaust gas flow expands through theturbine section to drive the compressor and the fan section. Thecompressor section typically includes low and high pressure compressors,and the turbine section includes low and high pressure turbines.

Internally cooling turbine blades may allow for use of higher gastemperatures, which may improve the engine's performance. Serpentinecore cooling passages have been used to cool turbine blades. Theserpentine cooling passage is arranged between the leading and trailingedge core cooling passages in a chord-wise direction. One typicalserpentine configuration provides “up” passages arranged near theleading and trailing edges fluidly joined by a “down” passage.

SUMMARY OF THE INVENTION

In one exemplary embodiment, an airfoil includes pressure and suctionside walls that extend in a chord-wise direction between leading andtrailing edges. The pressure and suction side walls also extend in aradial direction to provide an exterior airfoil surface. A coolingpassage is arranged between the pressure and suction walls. The coolingpassage has a first width in the chord-wise direction near the suctionside wall. A second width is in the chord-wise direction near thepressure side wall. A third width is between the pressure and suctionside walls. The third width is smaller than the first and second widths.

In a further embodiment of any of the above, a second cooling passage isarranged adjacent the cooling passage. The second cooling passage has akite shape.

In a further embodiment of any of the above, a plurality of aperturesconnect the second cooling passage with the cooling passage.

In a further embodiment of any of the above, the cooling passage has anhourglass shape.

In a further embodiment of any of the above, the third width is betweenabout 0.002 and about 0.100 inches (0.0508-2.540 mm).

In a further embodiment of any of the above, the first width defines asuction side cavity. The second width defines a pressure side cavity.The third width defines a passage between the pressure and suction sidecavities.

In a further embodiment of any of the above, the pressure and suctionside cavities are defined by a first internal wall and a second internalwall. Each of the first and second internal walls have a bent portion.

In a further embodiment of any of the above, the cooling passage is oneof a plurality of cooling passages. Some of the plurality of coolingpassages are in communication with one another in a serpentine coolingconfiguration.

In a further embodiment of any of the above, the cooling passage is oneof a plurality of cooling passages. Each of the plurality of coolingpassages receive air from a source and direct the air flow radially.

In another exemplary embodiment, a gas turbine engine includes acombustor section arranged fluidly between compressor and turbinesections. An airfoil is arranged in the turbine section. The airfoil haspressure and suction side walls that extend in a chord-wise directionbetween leading and trailing edges. The pressure and suction side wallsalso extend in a radial direction to provide an exterior airfoilsurface. A cooling passage is arranged between the pressure and suctionwalls. The cooling passage has a first width in the chord-wise directionnear the suction side wall. A second width is in the chord-wisedirection near the pressure side wall. A third width is between thepressure and suction side walls. The third width is smaller than thefirst and second widths.

In a further embodiment of any of the above, a second cooling passage isarranged adjacent the cooling passage. The second cooling passage has akite shape.

In a further embodiment of any of the above, a plurality of aperturesconnect the second cooling passage with the cooling passage.

In a further embodiment of any of the above, the cooling passage has anhourglass shape.

In a further embodiment of any of the above, the third width is betweenabout 0.002 and about 0.100 inches (0.0508-2.540 mm).

In a further embodiment of any of the above, the first width defines asuction side cavity. The second width defines a pressure side cavity.The third width defines a passage between the pressure and suction sidecavities.

In a further embodiment of any of the above, the pressure and suctionside cavities are defined by a first internal wall and a second internalwall. Each of the first and second internal walls have a bent portion.

In a further embodiment of any of the above, the cooling passage is oneof a plurality of cooling passages. Some of the plurality of coolingpassages are in communication with one another in a serpentine coolingconfiguration.

In a further embodiment of any of the above, the cooling passage is oneof a plurality of cooling passages. Each of the plurality of coolingpassages receives air from a source and directs the air flow radially.

In another exemplary embodiment, a core structure for a providing acooling passage in a gas turbine engine component includes a core bodythat has a first width in a chord-wise direction near a first wall. Asecond width is in the chord-wise direction near a second wall. A thirdwidth is in the chord-wise direction between the first and second walls.The third width is smaller than the first and second widths to form anhourglass shape.

In a further embodiment of any of the above, the first, second, andthird widths are formed from a die having rib and a pocket that receivesthe rib.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 schematically illustrates an example gas turbine engine.

FIG. 2 schematically illustrates an example engine section including afixed stage and a rotating stage.

FIG. 3 illustrates a perspective view of an airfoil having an internalcooling passage.

FIG. 4 illustrates a plan view of the airfoil illustrating directionalreferences

FIG. 5 illustrates a cross-sectional view of an airfoil taken along line5-5 of FIG. 3 .

FIG. 6 illustrates a perspective view of an airfoil having an exampleinternal cooling configuration.

FIG. 7 illustrates an example method step of manufacturing an airfoilcore.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. The fan section 22 drivesair along a bypass flow path B in a bypass duct defined within a housing15 such as a fan case or nacelle, and also drives air along a core flowpath C for compression and communication into the combustor section 26then expansion through the turbine section 28. Although depicted as atwo-spool turbofan gas turbine engine in the disclosed non-limitingembodiment, it should be understood that the concepts described hereinare not limited to use with two-spool turbofans as the teachings may beapplied to other types of turbine engines including three-spoolarchitectures.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects, a first (or low) pressure compressor 44 and a first (orlow) pressure turbine 46. The inner shaft 40 is connected to the fan 42through a speed change mechanism, which in exemplary gas turbine engine20 is illustrated as a geared architecture 48 to drive a fan 42 at alower speed than the low speed spool 30. The high speed spool 32includes an outer shaft 50 that interconnects a second (or high)pressure compressor 52 and a second (or high) pressure turbine 54. Acombustor 56 is arranged in exemplary gas turbine 20 between the highpressure compressor 52 and the high pressure turbine 54. A mid-turbineframe 57 of the engine static structure 36 may be arranged generallybetween the high pressure turbine 54 and the low pressure turbine 46.The mid-turbine frame 57 further supports bearing systems 38 in theturbine section 28. The inner shaft 40 and the outer shaft 50 areconcentric and rotate via bearing systems 38 about the engine centrallongitudinal axis A which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and fan drive gear system 48 may be varied. For example,gear system 48 may be located aft of the low pressure compressor, or aftof the combustor section 26 or even aft of turbine section 28, and fan42 may be positioned forward or aft of the location of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about five. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five 5:1. Low pressure turbine 46 pressure ratio is pressuremeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle. The geared architecture 48 may be an epicycle geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3:1 and less than about 5:1. Itshould be understood, however, that the above parameters are onlyexemplary of one embodiment of a geared architecture engine and that thepresent invention is applicable to other gas turbine engines includingdirect drive turbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet (10,668 meters). The flight condition of 0.8 Mach and35,000 ft (10,668 meters), with the engine at its best fuelconsumption—also known as “bucket cruise Thrust Specific FuelConsumption (‘TSFC’)”—is the industry standard parameter of lbm of fuelbeing burned divided by lbf of thrust the engine produces at thatminimum point. “Low fan pressure ratio” is the pressure ratio across thefan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The lowfan pressure ratio as disclosed herein according to one non-limitingembodiment is less than about 1.45. “Low corrected fan tip speed” is theactual fan tip speed in ft/sec divided by an industry standardtemperature correction of [(Tram ° R)/(518.7° R)]^(0.5). The “Lowcorrected fan tip speed” as disclosed herein according to onenon-limiting embodiment is less than about 1150 ft/second (350.5meters/second).

FIG. 2 shows a portion of an example engine, for example, a turbinesection 28. It should be understood, however, that this disclosure mayalso be provided in a compressor section. The section 28 includes afixed stage 60 that provides a circumferential array of vanes 63arranged axially adjacent to a rotating stage 62 with a circumferentialarray of blades 69. In the example, the vanes 63 include an outerdiameter portion 64 having hooks 65 that support the array of vanes 63with respect to a case structure. The vanes 68 may be secured to astructure, such as engine static structure 36, and remain stationaryrelative to the engine axis A. An airfoil 68 extends radially from theouter platform 64 to an inner diameter portion or platform 66. It shouldbe understood that the disclosed vane arrangement could be used for vanestructures cantilevered at the inner diameter portion of the airfoil. Aroot 74 of each blade 69 may be coupled to a rotor disk 72. The rotordisk 72 facilitates rotation of the blades 69 about the engine axis A.

Referring to FIG. 3 , a root 74 of each blade 69 is mounted to a rotordisk. The blade 69 includes a platform 76, which provides the inner flowpath, supported by the root 74. An airfoil 78 extends in a radialdirection R from the platform 76 to a tip 80. It should be understoodthat the turbine blades may be integrally formed with the rotor suchthat the roots are eliminated. In such a configuration, the platform 76is provided by the outer diameter of the rotor. The airfoil 78 providesleading and trailing edges 82, 84. The tip 80 is arranged adjacent to ablade outer air seal (not shown).

The airfoil 78 of FIG. 4 somewhat schematically illustrates an exteriorairfoil surface 79 extending in a chord-wise direction C from a leadingedge 82 to a trailing edge 84. The airfoil 78 is provided betweenpressure (typically concave) and suction (typically convex) walls 86, 88in an airfoil thickness direction T, which is generally perpendicular tothe chord-wise direction C. Multiple turbine blades 69 are arrangedcircumferentially in a circumferential direction A. The airfoil 78extends from the platform 76 in the radial direction R, or spanwise, tothe tip 80.

In the example, the airfoil 78 includes a serpentine cooling passage 90provided between the pressure and suction walls 86, 88. The serpentinecooling passage 90 provides a core cooling passage. The disclosedcooling passage arrangement may be used with other cooling passageconfigurations, including non-serpentine cooling passage arrangements.The cooling passage 90 receives cooling air from a cooling source 70 forinternally cooling the airfoil 78. The cooling source 70 may be a sourceradially inward of the airfoil 78, in some examples.

FIG. 5 illustrates an example cross-section of the airfoil 78 takenalong line 5-5 of FIG. 3 . The cooling passage 90 may include passageshaving different shapes, such as passage 106. The cooling passage 106 isarranged between hot outer walls 87, 89 of the pressure side 86 andsuction side 88, respectively. The cooling passage 106 has an“hourglass” shape. That is, the cooling passage 106 has a first width W₁near the suction side 88, a second width W₂ near the pressure side 86,and a third width W₃ between the first and second widths. The thirdwidth W₃ is smaller than the first and second widths W₁, W₂. In someexamples, the second width W₂ is smaller than the first width W₁. Thewidths W₁, W₂, W₃ are measured generally in the chordwise direction.

The third width W₃ generally divides the cooling passage 106 into apressure side cavity and a suction side cavity. The first width W₁ isthe widest point in the suction side cavity, while the second width W₂is the widest point in the pressure side cavity. The third width W₃provides a passage between the pressure and suction side cavities andcan be sized to meter the flow between the pressure and suction sidecavities. The third width W₃ may be at least about 0.002 inches (0.0508mm), for example. In an example, the third width W₃ may be between about0.002 and 0.100 inches (0.0508-2.540 mm). In a further example, thethird width W₃ is about 0.030 inches (0.762 mm).

The widths W₁ and W₂ provide a large surface along the hot outer wallsfor cooling of the airfoil 78. The width W₃ is near a camber line of theairfoil. This gives the cooling passage 106 a smaller cross-sectionalarea, which may provide more efficient cooling.

The first and second widths W₁, W₂ are defined by a first internal wall110 and a second internal wall 118. Each of the walls 110, 118 has abend 116, which defines the third width W₃. The bend 116 divides eachwall 110, 118 into a first portion 112 and a second portion 114 arrangedat an angle Θ relative to the first portion 112. The angle Θ may begreater than 90°, for example. The angle Θ for each wall 110, 118 may bethe same or may be different. The walls 110, 118 are connected to thehot outer walls 87, 89, and are not connected to one another. The bend116 permits the walls 110, 118 to flex and move with the outer walls 87,89, when the outer walls 87, 89 expand axially. In other words, as theouter walls 87, 89 expand, the first width W1, second width W2, andthird width W3 increase. The bend 116 allows the walls 110, 118 to flexso that each of the widths W1, W2, and W3 can increase a differentamount. This may reduce stresses in the airfoil 78 as the outer wallsexpand.

The airfoil 78 may include multiple “hourglass” shaped passages, such aspassages 107, 109. In this example, the airfoil 78 includes several kiteor diamond shaped passages 105, 108, 111, 113 adjacent the hourglasspassages. The diamond shaped passages 105, 108, 111, 113 are sandwichedbetween the hourglass passages 106, 107, 109. The diamond passages 105,108, 111, 113 are isolated from the hot outer walls, and thus requireless cooling air in those passages. In some examples, the diamondpassages 105, 108, 111, 113 may transport cold air to other locations ofthe airfoil 78 or other components in the engine. In another example,the diamond passages 105, 108, 111, 113 are plenums with a small amountof cooling air.

FIG. 6 illustrates an example airfoil 78. A leading edge cooling passage126 is arranged near the leading edge 82, while a trailing edge coolingpassage 124 is arranged near the trailing edge 84. A plurality ofhourglass shaped cooling passages and diamond shaped cooling passagesare arranged between the leading and trailing edge cooling passages 122,124. In the illustrated example, the airfoil 78 includes three hourglassshaped cooling passages 106, 107, 109, and four diamond shaped coolingpassages 105, 108, 111, 113. In other examples, more or fewer coolingpassages may be utilized.

In one example, the cooling passages 106, 107, 109 are arranged in aserpentine configuration. In this example, the cooling passage 106 isjoined to the cooling passage 107 at a radially outer bend, and thecooling passage 107 is joined to the cooling passage 109 at a radiallyinner bend. The cooling passage 106 receives air from a cooling source70 (shown in FIG. 3 ), the cooling air travels radially outward throughthe cooling passage 106, then radially inward through cooling passage107, and radially outward through cooling passage 109. The diamondpassages 105, 108, 111, 113 may all flow radially outward in thisexample.

In another example, the cooling passages 106, 107, 109 are arranged witha radial flow design. In this example, cooling air moves radiallyoutward through each of the cooling passages 106, 107, 109. Cooling airmay also move radially outward through cooling passages 126, 105, 108,111, 113, and 124. The cooling air from the trailing edge coolingpassage 124 may also flow axially out cooling edge slots. In anotherexample, the cooling air moves radially inward through each of thecooling passages. This arrangement may be useful when the airfoil 78 isa vane, for example.

In some examples, the cooling passages may include trip strips,deptowarts, dimples, and/or pin fins on the internal surface of the hotwalls 87, 89. The hot walls 87, 89 may have film cooling holes 121 forcommunicating cooling air to an exterior surface of the airfoil 78. Theleading edge passage 126 may have film cooling holes 122 incommunication with an exterior surface of the leading edge 82. In someembodiments, cooling apertures 120 connect adjacent cooling passages.The cooling apertures 120 may be holes, slots, gaps, or have anothergeometry. For example, cooling apertures 120 may connect a diamondshaped cooling passage, such as cooling passage 108 or 111, with an“hourglass” cooling passage, such as cooling passage 106 or 107. Thesecooling apertures 120 may provide impingement cooling in the coolingpassages.

The disclosed airfoil 78 may be cast or additively manufactured. FIG. 7illustrates an example method step of manufacturing a core for castingthe example airfoil 78. In this example, the airfoil core is cast in acore die having a first block 150 and a second block 152. Each of thefirst and second blocks 150, 152 has a plurality of pockets 154. Theinternal walls of the airfoil, such as bent walls 110, 118 are formedusing ribs 156, 158. Ribs 158 are aligned with the separation directionof blocks 150 and 152, whereas ribs 156 are not. After the airfoil coreis cast in the dies about the ribs 156, 158, to prevent ribs 156 frombeing trapped by the core during block separation, the ribs 156 are slidoutward relative to the airfoil core in the direction of ribs 156 andinto the pockets 154. The first and second blocks 150, 152 may beseparated in a direction aligned with ribs 158 to release the airfoilcore. The first and second blocks 150, 152 may be a metallic material,for example. The core may be formed by injecting a ceramic material intothe core die, for example. The core may then be used to cast an airfoil.

In another example, the airfoil core may be manufactured usingsacrificial inserts. In this example, a sacrificial insert is placed inthe core die to form the walls 110, 118. The sacrificial insert may be athermoplastic piece, for example. After the die is injected withmaterial, the sacrificial insert is melted out, leaving a core havingthe above described internal features. The core may be used in a wax dieto manufacture the airfoil, for example. In some examples, the die isinjected with a ceramic material. After the core die is removed, theceramic is fired to melt the sacrificial insert. The core may then beused to cast an airfoil.

In another example, the airfoil may be additively manufactured. Theairfoil may be formed by sintering a metal powder one layer at a time toform the airfoil and internal walls. In one example, the airfoil isadditively manufactured starting at the inner diameter portion andmoving toward the outer diameter, relative to the engine axis A.

Known airfoil cooling arrangements have inner walls that are cold andstiff, and outer walls that are hot, causing higher thermal stress inthe component. The disclosed cooling passage arrangement may reduce therequired cooling flow to the airfoil by reducing the cross-sectionalareas of the cooling cavities near the airfoil walls without increasingthermal stresses. The pressure side and suction side cavities are formedby bent walls that are not attached to one another, forming a passagebetween the pressure and suction side cavities. That is, the walls areseparate from one another. This arrangement allows the ribs to movefreely with the hot expanding outer walls, reducing thermal stress.

In this disclosure, “generally axially” means a direction having avector component in the axial direction that is greater than a vectorcomponent in the circumferential direction, “generally radially” means adirection having a vector component in the radial direction that isgreater than a vector component in the axial direction and “generallycircumferentially” means a direction having a vector component in thecircumferential direction that is greater than a vector component in theaxial direction.

Although the different examples have specific components shown in theillustrations, embodiments of this invention are not limited to thoseparticular combinations. It is possible to use some of the components orfeatures from one of the examples in combination with features orcomponents from another one of the examples.

Although an embodiment of this invention has been disclosed, a worker ofordinary skill in this art would recognize that certain modificationswould come within the scope of this disclosure. For that reason, thefollowing claims should be studied to determine the true scope andcontent of this disclosure.

The invention claimed is:
 1. A method of manufacturing a core structurefor casting an airfoil, comprising: providing a core die having a firstblock and a second block, and a plurality of ribs extending from thefirst block and the second block; casting an airfoil core in the coredie about the plurality of ribs, the plurality of ribs arranged to forma first, second, and third cooling passage core, the first and thirdcooling passage cores having an hourglass shape and the second coolingpassage core arranged between the first and third cooling passage coresand having a kite shape, wherein a perimeter of the second coolingpassage core includes an opposed pair of acute angles and an opposedpair of obtuse angles that establish the kite shape; and separating thefirst block and the second block to release the airfoil core from thecore die.
 2. The method of claim 1, wherein the first and second blockseach have a plurality of pockets aligned with the ribs.
 3. The method ofclaim 2, comprising sliding the ribs into the pockets before separatingthe first and second blocks.
 4. The method of claim 3, wherein: the stepof sliding the ribs into the pockets occurs subsequent to the step ofcasting the airfoil core in the core die, but prior to the step ofseparating the first and second blocks.
 5. The method of claim 1,wherein at least one of the plurality of ribs extends from the firstblock in a first direction, and at least one of the plurality of ribsextends from the second block in a second direction.
 6. The method ofclaim 5, wherein the separating step comprises moving the first block ina direction opposite the first direction and moving the second block ina direction opposite the second direction.
 7. The method of claim 1,wherein the plurality of ribs are formed from a sacrificial materialthat is configured to melt out of the airfoil core.
 8. The method ofclaim 1, wherein the first and second blocks are formed from a metallicmaterial.
 9. The method of claim 1, wherein the airfoil core is formedfrom a ceramic material.
 10. The method of claim 1, wherein: theplurality of ribs includes a first set of ribs and a second set of ribs;and the step of providing the core die occurs such that the first set ofribs extend outwardly from the first block towards the second block andsuch that the second set of ribs extend outwardly from the second blocktowards the first block to abut a respective rib of the first set ofribs prior to the casting step.
 11. The method of claim 9, wherein: thefirst set of ribs and the second set of ribs establish a first corecavity, a second core cavity and a third core cavity; the forming stepincludes forming the first cooling passage core in the first corecavity, forming the second cooling passage core in the second corecavity, and forming the third cooling passage core in the third corecavity; a minimum width of the first core cavity is established at aposition in which the first set of ribs abut the second set of ribs; aminimum width of the third core cavity is established at a position inwhich the first set of ribs abut the second set of ribs; and a maximumwidth of the second core cavity is established at a position in whichthe first set of ribs abut the second set of ribs.
 12. The method ofclaim 1, wherein: the step of providing the core die occurs such thatthe first set of ribs are spaced apart from the second block and suchthat the second set of ribs are spaced apart from the first block. 13.The method of claim 12, wherein: the first block includes a first set ofpockets in a thickness of the first block, the first set of pocketsaligned with respective ribs of the first set of ribs; and the secondblock includes a second set of pockets in a thickness of the secondblock, the second set of pockets aligned with respective ribs of thesecond set of ribs.
 14. The method of claim 13, comprising: sliding thefirst set of ribs into the respective pockets of the first set ofpockets subsequent to the step of casting the airfoil core in the coredie, but prior to the step of separating the first and second blocks;and sliding the second set of ribs into the respective pockets of thesecond set of pockets subsequent to the step of casting the airfoil corein the core die, but prior to the step of separating the first andsecond blocks.
 15. The method of claim 14, wherein: the first coolingpassage core and the second passage core are established on oppositesides of one of the ribs of the first set of ribs and on opposite sidesof one of the ribs of the second sets of ribs that abuts the one of theribs of the first set of ribs; and the second passage core and the thirdpassage core are established on opposite sides of another one of theribs of the first set of ribs and on opposite sides of another one ofthe ribs of the second sets of ribs that abuts the another one of theribs of the first set of ribs.